Gas turbine engines are well-known for providing propulsion and power production. Compared to other engine types, a gas turbine engine has the advantage that its rotors undergo purely rotational motion, and it can therefore operate at high speed with minimum vibration. In such engines, the theory of operation is described thermodynamically by the Brayton cycle: air is compressed isentropically, combustion of air/fuel mixture occurs, and expansion over turbine blades occurs isentropically back to the starting pressure. Such engines, however, typically operate efficiently only within a relatively narrow band of engine speeds.
Conventional modern turbine engines have to run at high temperatures to produce enough work to drive multiple stages of compression to achieve high pressure ratios and high efficiency. The high temperatures and pressures drive up design complexity and life cycle costs (development, production, maintenance). Also, these turbine engines are designed for optimum performance at one design point which causes the engines to operate less efficiently during off design conditions. That is, conventional turbine engines cannot produce the optimum higher pressure ratios at lower cruise power settings.
Another limitation of conventional turbine engines is the use of a rotor for each compressor and turbine stage, as well as a connecting shaft linking the compressor and turbine, which significantly increases the engine length, weight, and cost. Also, a limiting factor is the ability of steel, nickel, ceramic, or other engine materials to withstand extreme heat and pressure. These extreme temperatures require elaborate/complex secondary flow cooling circuits to maintain acceptable material properties, especially at the high pressure turbine bores. Unfortunately, these cooling systems reduce engine performance and add undesirable weight. Even when cooling systems are used in turbine engines, there is centrifugal and thermal growth of the rotors. Maintaining a gap between the rotor tips and the engine shroud to account for thermal growth causes large tip leakage losses and lower component performance and efficiency. Conventional turbine engines also have complex bearing systems that operate near maximum temperature limits and endure extreme shaft dynamics.
There are a myriad of known gas turbine engine configurations. One early example is illustrated in FIG. 1. The engine in FIG. 1 was designed by Hans von Ohain in 1937 and was designated the He.S3 turbojet engine. The combustor of the Ohain engine is positioned in the large unused space in front of the radial-flow compressor. Airflow through the He.S3 engine followed a generally S-shaped configuration.
U.S. Pat. No. 2,694,291 to Rosengart describes a rotor and combustion chamber arrangement for gas turbines. The turbine includes a stationary combustion chamber that is generally toroidal shaped and has a continuous opening at its inner periphery. Mounted on the rotor are hollow blades designed so that air and exhaust gases pass between the blades while cooling air passes within the hollows of the blades to facilitate cooling.
Another gas turbine example is described in U.S. Pat. No. 3,269,120 to Sabatiuk. Sabatiuk discloses a gas turbine engine having compressor and turbine passages in a single rotor element. The engine has axial flow compressor passages and radial flow turbine passages in a single rotor. Flow through the compressor passages is in a direction generally parallel to the axis of the rotor element, and flow through the turbine passages is in a radial direction at least for a portion of the length of the passages.
U.S. Pat. No. 3,892,069 to Hansford describes a propulsion unit for an aircraft that includes rotor means incorporating a multi-bladed fan which over an outer peripheral region thereof defines centrifugal flow compressor passages and centripetal flow turbine passages and an annular combustor encircling the rotor means. The combustor has inlet means for directing air from the compressor passages into the combustor and outlet means for directing combustion gases from the combustor into the turbine passages to drive the rotor means. Hansford's propulsion unit includes an air intake leading to a series of circularly distributed centrifugal flow compressor passages and includes an annular combustor which is of substantially toroidal shape and defines a combustion chamber of part circular cross-section.
In addition, Dev in U.S. Pat. No. 6,988,357 discloses a gas turbine engine including a combustion chamber section, a turbine section, and a compressor section. The turbine section surrounds the combustion chamber, and the compressor section surrounds the turbine section. The rear rotor of the turbine engine includes an integral compressor section on the outside and a turbine section on the inside.
Given the limitations of conventional gas turbine engines, there is a need for an improved engine that minimizes weight and fuel consumption while maximizing thrust and efficiency.